Pressure seal with non-metallic wear surfaces

ABSTRACT

A gas turbine engine has a first component and a second component. The first and second components have a high-pressure chamber on one side and a low pressure chamber on an opposed side. A three sided seal has one side facing each of the first and second components, and a third side facing a third component. At least one non-metallic wear surface is between one of the three sides of the seal and the facing component.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.13/688,340, filed Nov. 29, 2012.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

This invention was made with government support under Contract No.N00019-02-C-3003 awarded by the United States Navy. The Government hascertain rights in this invention.

BACKGROUND

This application relates to a pressure seal having three surfaces insliding contact, with at least one of the surfaces being provided with anon-metallic wear surface.

Gas turbine engines are known, and typically include a fan deliveringair into a compressor. The air is compressed in the compressor, anddelivered into a combustion section. The air is mixed with fuel in thecombustion section and ignited. Products of the combustion passdownstream over turbine rotors, driving them to rotate.

There are a number of areas within a gas turbine engine where a highpressure chamber must be maintained separate from a low pressurechamber. Thus, various locations require a reliable seal.

One type of seal is a U or J-shaped seal having three sides. Twocomponents contact two of the sides of the seal, and some form ofretention member may contact a third side. The seal is generallymoveable along the three surfaces.

In the past, seals in this particular application have always beenformed of metallic materials. This has led to undue wear.

While non-metallic seals have been proposed for many applications, theyhave not been proposed in gas turbine engines between high and lowpressure chambers, where a three sided sealing application is used.

SUMMARY

In a featured embodiment, a gas turbine engine has a first component anda second component. The first and second components have a high-pressurechamber on one side and a low pressure chamber on an opposed side. Athree sided seal has one side facing each of the first and secondcomponents, and a third side facing a third component. At least onenon-metallic wear surface is between one of the three sides of the sealand the facing component.

In another embodiment according to the previous embodiment, there arenon-metallic wear surfaces between each of the three sides of the sealand the facing component.

In another embodiment according to any of the previous embodiments, oneof the components is a liner segment of an exhaust system in the gasturbine engine.

In another embodiment according to any of the previous embodiments,another the component is a moving liner segment.

In another embodiment according to any of the previous embodiments, thethird component is a retention element for positioning the seal.

In another embodiment according to any of the previous embodiments, oneof the components is a case segment for the gas turbine engine.

In another embodiment according to any of the previous embodiments, oneof the components is a duct segment for the gas turbine engine.

In another embodiment according to any of the previous embodiments, thewear surface is on the seal.

In another embodiment according to any of the previous embodiments, thewear surface is on at least one of the components.

In another embodiment according to any of the previous embodiments, thethird side may sometimes be in contact with the third component, andsometimes be spaced from the third component during differentoperational points in the operation of the gas turbine engine.

In another embodiment according to any of the previous embodiments, theretention segment has a pair of retention elements on opposed positionfaces of one of the three sides of the seal.

In another embodiment according to any of the previous embodiments, thethree sided seal floats relative to each of the first, second and thirdcomponents such that it is movable relative to each of the components.

These and other features of this application will be best understoodfrom the following specification and drawings, the following of which isa brief description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 shows a location in the FIG. 1 gas turbine engine receiving aseal.

FIG. 3A shows a first seal embodiment.

FIG. 3B shows another seal embodiment.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flowpath B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flowpath C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a turbofan gas turbine engine in the disclosednon-limiting embodiment, it should be understood that the conceptsdescribed herein are not limited to use with turbofans as the teachingsmay be applied to other types of turbine engines including three-spoolarchitectures.

The engine 20 generally includes a low speed spool 30 and a high speedspool 32 mounted for rotation about an engine central longitudinal axisA relative to an engine static structure 36 via several bearing systems38. It should be understood that various bearing systems 38 at variouslocations may alternatively or additionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and high pressure turbine54. A combustor 56 is arranged between the high pressure compressor 52and the high pressure turbine 54. A mid-turbine frame 57 of the enginestatic structure 36 is arranged generally between the high pressureturbine 54 and the low pressure turbine 46. The mid-turbine frame 57further supports bearing systems 38 in the turbine section 28. The innershaft 40 and the outer shaft 50 are concentric and rotate via bearingsystems 38 about the engine central longitudinal axis A which iscollinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a planetary gearsystem or other gear system, with a gear reduction ratio of greater thanabout 2.3 and the low pressure turbine 46 has a pressure ratio that isgreater than about 5,In one disclosed embodiment, the engine 20 bypassratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout 5:1,Low pressure turbine 46 pressure ratio is pressure measuredprior to inlet of low pressure turbine 46 as related to the pressure atthe outlet of the low pressure turbine 46 prior to an exhaust nozzle.The geared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.5:1. It should be understood, however, that theabove parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present invention is applicable toother gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition —typically cruise at about 0.8 Machand about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft,with the engine at its best fuel consumption —also known as “bucketcruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industrystandard parameter of lbm (pounds mass) of fuel being burned divided bylbf (pounds force) of thrust the engine produces at that minimum point.“Low fan pressure ratio” is the pressure ratio across the fan bladealone, without a Fan Exit Guide Vane (“FEGV”) system. The low fanpressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram °R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second.

FIG. 2 shows a location within a gas turbine engine, which may be in anozzle portion of the gas turbine engine, downstream of the turbinesection 28. A high pressure area 70 is separated from a lower pressurearea 72 by a seal 74. Seal 74 has three sides 77, 79 and 81, all ofwhich may be in contact with a seal surface on an associated component76, 78 and 80. As shown, the sides 77, 79 and 81 may be also out ofcontact with a component 77, 78 and 80 but each face a component.

Component 78 may be a static liner segment of an exhaust system, and hasa sealing surface 82 positioned to be in contact with the surface 79 onthe seal 74. A pair of retention elements 76 face, and may contact, thesurface 77 on the seal 74. As shown in FIG. 2, the retention elements 76may also be spaced away from the surface 77 on the seal 74 under much ofits operational life. Another component 80 may be a moving linersegment, and have an end surface in contact with a surface 81 on theseal 74.

In other embodiments, one of the components may be a case segment for agas turbine engine, or a duct segment for a gas turbine.

FIG. 3A is an enlarged view of the seal embodiment.

As shown in FIG. 3A, the seal floats between the three surfaces, and asit moves there is wear. Thus, the provision of a non-metallic wearsurface 82 between the surfaces 78 and 79 provides longer life for theseal and the components.

In embodiments, wear surfaces may be applied on each of the components76, 80, and 78. The wear surfaces may be polytetraflouroethylene (PTFE)or thermoplastic polymer or other appropriate non-metallic materials.The non-metallic surfaces may be formed of carbon, silicon, ceramic orother composite-based materials and may be selected due to wear andlubrication characteristics. The materials may be applied as pads,segmented strips, or continuous strips.

FIG. 3B shows another embodiment 174. In embodiment 174, the three sides176, 178 and 180 of the seal receive non-metallic pads 177, 179 and 181,respectively.

The pads or other non-metallic materials may be mechanically secured tothe seal or other components via rivets, fasteners, and etcetera.

Further, the entire seal may be formed of non-metallic materials, oralternatively, the members 76, 78 and 80 may be formed of non-metallicmaterials. That is, there need not be a non-metallic material secured tothe underlying substrate, but rather the substrate itself may be formedof the non-metallic materials.

Although embodiments of this invention have been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this invention. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this invention.

The invention claimed is:
 1. A gas turbine engine comprising: a firstcomponent and a second component, said first and second componentshaving a high-pressure chamber on one side, and a low pressure chamberon an opposed side, and a three sided seal having one side facing eachof said first and second components, and having a third side facing athird component, and there being at least one non-metallic wear surfacebetween one of said three sides of said seal and said facing component;and said three sided seal floats relative to each of said first, secondand third components such that it is movable relative to each of saidcomponents.
 2. The gas turbine engine as set forth in claim 1, whereinthere are non-metallic wear surfaces between each of the three sides ofsaid seal and said facing component.
 3. The gas turbine engine as setforth in claim 2, wherein one of said components is a liner segment ofan exhaust system in the gas turbine engine, another said component is amoving liner segment, and said third component is a retention elementfor positioning the seal.
 4. The gas turbine engine as set forth inclaim 3, wherein said retention segment has a pair of retention elementson opposed position faces of one of said three sides of said seal. 5.The gas turbine engine as set forth in claim 1, wherein one of saidcomponents is a case segment for the gas turbine engine.
 6. The gasturbine engine as set forth in claim 1, wherein one of said componentsis a duct segment for the gas turbine engine.
 7. The gas turbine engineas set forth in claim 1, wherein said wear surface is on the seal. 8.The gas turbine engine as set forth in claim 1, wherein said wearsurface is on at least one of the components.
 9. The gas turbine engineas set forth in claim 1, wherein said third side may sometimes be incontact with said third component, and sometimes be spaced from saidthird component during different operational points in the operation ofthe gas turbine engine.